At present, efforts are being undertaken to develop hypersonic or reusable space vehicles capable of reaching speeds as high as Mach 12. Examples of such vehicles include, for example, missiles, hypersonic cruise vehicles, and spacecraft such as the space shuttle.
Such hypersonic or reusable space vehicles are, of course, subject to extreme temperature fluctuations within the vehicle's envelope of performance. Specifically, the leading edges, flight control surfaces and a substantial portion of the external surfaces of such vehicle support structures, or frames, as well as the internal construction associated with engines necessary to power the vehicle require that thermal design parameters incorporate means for ensuring structural survivability during short periods of high heat flux. Thermal protection systems for hypersonic vehicles essentially fall into two categories: insulative and ablative. Insulative systems such as those used on the space shuttle have two advantages: (i) they are generally lighter in weight than ablative systems and (ii) they maintain a constant outer vehicle surface, whereas with ablative systems, recession of the outer surface occurs thus changing the aerodynamic shape of the vehicle. However, existing insulative systems are limited in the maximum allowable temperature (or heat flux) at the outer surface (mostly below ˜1600 deg. C.), whereas ablative systems can be used to much higher temperatures (and heat fluxes). There exists a need to provide adequate thermal protection to hypersonic or reusable space vehicles in the event of a high heat load event that combines the most desirable attributes of both the insulative and ablative thermal protection systems. Such a system ideally also realizes other positive attributes such as cost and weight reduction.